1. Technical Field
The present invention relates to gas turbine engines in general, and to blade tip clearance therein in particular.
2. Background Information
In an axial flow gas turbine engine, air is compressed in a compressor section, mixed with fuel and combusted in a combustor section, and expanded through a turbine section that, via one or more shafts, drives the compressor section. The overall efficiency of such engines is a function of, among other factors, the efficiency with which the compressor section compresses the air. The compressor section typically includes a low pressure compressor driven by a shaft connected to a low pressure turbine in the turbine section, and a high pressure compressor driven by a shaft connected to a high pressure turbine in the turbine section. The high and low compressors are multi-stage, wherein the air flows in the axial direction through a series of rotors and stators that are concentric with the axis of rotation (longitudinal axis).
The stages are arranged in series. Each stage compresses the air passing through the engine and thereby incrementally increases the pressure of the air. The total pressure increase through the compressor is the sum of the incremental pressure increases through each stage, adjusted for any flow losses. Thus, in order to maximize the efficiency of the gas turbine engine, it would be desirable, at a given fuel flow, to maximize the pressure rise (hereinafter referred to as “pressure ratio”) across each stage of the compressor.
For a variety of reasons, including efficiency, it is highly desirable to minimize the clearance between the blade tips of a rotor and the casing surrounding the rotor. Prior art solutions to maintaining blade tip clearance include the use of abradables and active clearance control systems that manipulate the radial position of the casing surrounding the rotor. A problem with an abradable system is that it is not adjustable once the seal is abraded. Once set, the clearance depends solely on the thermal and centrifugal response of the rotor and the casing. A problem with prior art active clearance systems is their response time. Prior art active clearance systems often utilize the flow (or lack of flow) of cooling air as a mechanism to move the casing via thermal expansion or contraction and thereby achieve the desired clearance. Such systems are still subject to the casing s thermal response time, and the mismatch of the casing s thermal response to the rotor disc s thermal response.
FIG. 1 is a diagrammatic graph of blade tip clearance versus response time typical of a prior art gas turbine engine. At position A, the engine is running at a steady-state idle condition (i.e., low rpm s, low core gas temperature). At position B, the engine is sharply accelerated (e.g., acceleration for takeoff). As a result of the acceleration, the blade tip clearance decreases dramatically, reaching a minimum clearance at position C. At this point, the change in clearance is almost entirely attributable to mechanical growth of the rotor assembly as a result of the centrifugal loading on the rotor assembly and the thermal growth of the blade.
The increase in core gas temperature that accompanies the acceleration next causes thermal radial growth of the casing surrounding the rotor assembly (from position C to position D). The decrease in clearance between positions D and E is attributable to the eventual thermal growth of the rotor disk. The greater mass of the rotor assembly, in particular the rotor disk, causes it to have a slower thermal response than that of the casing. The eventual stabilized clearance at position E is a function of the final temperatures of the disc, blades, casing, the centrifugal pull, and the coefficient of thermal expansion of each part of the rotor assembly and casing.
The significant increase in clearance between positions E and F is a result of a deceleration. The deceleration causes a decrease in the centrifugal loading on the rotor assembly and rapid cooling of the blades, which results in a decrease in the mechanical growth of the rotor assembly. Under normal conditions, the clearance will decrease in the manner suggested by the line extending between positions F and I. Here again, however, the clearance will depend on the thermal responses of the disc and the casing.
If, however, a sharp acceleration is performed between positions G and H, the clearance in a prior art compressor will decrease significantly. The decrease in clearance results from the combination of: 1) the almost immediate mechanical growth of the rotor assembly; 2) the already decreased casing inner radial dimension as a result of the casing s relatively immediate thermal response rate following the deceleration; and 3) the lack of decrease in rotor assembly attributable to the rotor s relatively slow thermal response. With respect to the latter two factors, the difference in thermal response between the casing and the rotor assembly creates a situation where the casing has already substantially returned to its pre-thermal growth dimension, but the rotor assembly has not yet returned to its pre-thermal growth dimension. The resultant rub-out of the seal is shown at position H. In prior art gas turbine engines, the clearance at position H is often chosen as the worst-case clearance and blade tip clearance is designed to accommodate the operating scenario present at position H. As a result, the blade tip/casing clearances are minimum at position H, but less than optimum under normal operating conditions (e.g., positions A-G).
Thus, what is needed is an improved apparatus and method for maintaining a desired blade tip clearance throughout transient and steady-state operation of the gas turbine engine.